Typical gas turbine engine fuel supply systems include a fuel source, such as a fuel tank, and one or more pumps that draw fuel from the fuel tank and deliver pressurized fuel to the fuel manifolds in the engine combustor via a main supply line. The main supply line may include one or more valves in flow series between the pumps and the fuel manifolds. These valves generally include at least a metering valve and a pressurizing-and-shutoff valve downstream of the metering valve. In addition to the main supply line, many fuel supply systems may also include a bypass flow line connected upstream of the metering valve that bypasses a portion of the fuel flowing in the main supply line back to the inlet of the one or more pumps, via a bypass valve. The position of the bypass valve is controlled to maintain a substantially fixed differential pressure across the main metering valve.
Although the above-described fuel supply system configuration is generally safe and reliable, it can suffer certain drawbacks depending on the particular aircraft configuration. For example, these fuel systems can exhibit undesirable levels of fuel heating due to the continuous flow of fuel through the bypass valve and fuel pump. Commonly, the waste heat that is generated in an aircraft is cooled by the engine fuel supply system so that the heat can subsequently be extracted in the gas turbine engine. However, in many systems it is not possible to sink all of the waste heat into the fuel due to maximum temperature limits on the fuel and fuel wetting components. This may result in adding a heat exchanger in the engine fan ducts, which can introduce additional weight and noise, and can also reduce net engine thrust.
Hence, there is a need for a gas turbine engine fuel supply system that does not increase fuel temperature to undesirable levels during operation and/or does not rely on additional heat exchangers. The present invention addresses at least these needs.